Turbine blade with improved trailing edge cooling

ABSTRACT

A turbine blade ( 110 ) is provided, where an exterior surface of the turbine blade ( 110 ) is exposed to a hot combustion gas ( 125 ) above a flow path line ( 141 ). The turbine blade ( 110 ) includes a trailing edge pin bank cooling channel ( 118 ) in an airfoil section ( 112 ), a cooling air supply channel ( 120 ) in a root section ( 114 ), and a channel ( 122 ) shaped with a radius of curvature ( 126 ), to extend the channel ( 122 ) across the flow path line ( 141 ) and interconnect the cooling air supply channel ( 120 ) to the trailing edge pin bank cooling channel ( 118 ). A width of the trailing edge pin bank cooling channel ( 118 ) is adjusted, such that the width at an inner diameter region ( 128 ) and an outer diameter region ( 131 ) is less than the width at an intermediate region ( 130 ) between the inner and outer diameter regions ( 128,131 ).

FIELD OF THE INVENTION

The present invention relates to turbine blades, and more specifically,to a turbine blade having an improved cooling system.

BACKGROUND OF THE INVENTION

Typically, gas turbine engines include a compressor for compressing air,a combustor for mixing the compressed air with fuel and igniting themixture, and a turbine blade for producing power. Combustors oftenoperate at high temperatures that may exceed 2500 degrees Fahrenheit.Typical turbine combustor configurations expose turbine blade assembliesto these high temperatures. As a result, turbine blades often containcooling systems for prolonging the life of the blades and reducing thelikelihood of failure as a result of the exposure to the hightemperatures.

FIG. 1 illustrates a conventional turbine blade 10 including an airfoilsection 12 joined to a root section 14 at a platform section 16.Generally, the airfoil section 12 includes a pressure sidewall and asuction sidewall (not shown) joined along a leading edge 11 and atrailing edge 13 and extending radially outward from the platformsection 16 at an inner diameter region 28 to a tip 21 at an outerdiameter region 31. The airfoil section 12 includes a cooling system 19with a serpentine cooling passage that receives cooling fluid from aninlet in the root section 14. A cooling fluid flow 39 from the rootsection 14 is received in the serpentine cooling passage, which includesa first channel 44 and a second channel 46 joined by a first turn 45 atthe tip 21, and a third channel 48 joined to the second channel 46 by asecond turn 47 at the inner diameter region 28. During the gas turbineoperation, a hot combustion fluid flow 25 passes around an exterior ofthe airfoil section 12 above the platform section 16 which defines aflow path line 41. The flow path line 41 defines the radial extent ofthe area of the blade which is exposed to the hot combustion gas.Additionally, a cooling fluid flow 23 passes from a supply channel 20into a third channel 48, through parallel impingement orifices 51 and achannel 22 and then into an impingement trailing edge cooling channel 18along the trailing edge 13, after which the cooling fluid passes throughimpingement orifices 52 between spaced-apart ribs 58, before exiting theairfoil section 12 through the trailing edge 13.

FIG. 2 is a closer view of the channel 22 that connects the thirdchannel 48 in the airfoil section 12 to the impingement trailing edgecooling channel 18 in the airfoil section 12. As illustrated in FIG. 2,the channel 22 defines a cross-sectional flow area for the cooling fluidat an entrance to the impingement trailing edge cooling channel 18,between a rib 38 and a surface 27 of the airfoil section 12. The surface27 has a radius of curvature 26. As further illustrated in FIG. 2, thechannel 22 is positioned above the flow path line 41 and above theplatform section 16 of the turbine blade 10 such that the radius ofcurvature 26 is located entirely in the region of the blade exposed tothe hot combustion gas. In an exemplary embodiment, the channel 22 is acollection chamber through which the cooling fluid 23 passes at a slowor static rate.

FIG. 3 illustrates a conventional ceramic core 64 used to form theconventional turbine blade 10 of FIG. 2 during a casting process, asappreciated by one of skill in the art. The core 64 includes a ribopening 70 corresponding to the rib 38 of the turbine blade 10 and achannel portion 68 with a radius of curvature 26 that corresponds to thechannel 22 of the turbine blade 10 subsequent to the casting process.The design of the core 64, including the radial position of the ribopening 70 as well as the radius of curvature 26 of the channel portion68, are determinative of the cross-sectional flow area of the coolingfluid which will pass through the entrance to the impingement trailingedge cooling channel 18 of the turbine blade 10. The core 64 alsoincludes a channel portion 72 corresponding to the impingement trailingedge cooling channel 18 of the turbine blade 10.

FIG. 8 illustrates the impingement trailing edge cooling channel 18 ofthe conventional turbine blade 10. Reflecting the tapered width of theblade along its radial axis, the width 32 of the cooling channel 18(measured between the pressure side 15 and the suction side 17perpendicular to a radial axis of the blade) at the inner diameterregion 28 is greater than the width 34 of the cooling channel at theintermediate region 30, and the width 34 of the cooling channel 18 atthe intermediate region 30 is greater than the width 36 of the coolingchannel 18 at the outer diameter region 31. Thus, the width 32,34,36 ofthe cooling channel 18 continuously increases from the inner diameterregion 28 to the outer diameter region 31. The turbine blade 10 has anouter thickness 29 at the inner diameter region 28. As appreciated byone of skill in the art, the width of the channel 22, and thus the flowof cooling fluid through the channel 22 and into the cooling channel 18is responsive to the width of the cooling channel 18.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is explained in the following description in view of thedrawings that show:

FIG. 1 is a cross-sectional side view of a conventional turbine blade;

FIG. 2 is a partial view of the cross-sectional side view of FIG. 1 atan entrance to an impingement trailing edge cooling channel;

FIG. 3 is a cross-sectional side view of a conventional core;

FIG. 4 is a top view of a turbine blade in accordance with an embodimentof the present invention;

FIG. 5 is a cross-sectional side view of the turbine blade of FIG. 4along the line 5-5;

FIG. 6 is a cross-section side view of a core used to form the turbineblade of FIG. 5;

FIG. 7 is a cross-sectional end view of the turbine blade of FIG. 4along the line 7-7; and

FIG. 8 is a cross-sectional end view of the turbine blade of FIG. 1along the line 8-8

DETAILED DESCRIPTION OF THE INVENTION

The present inventors have recognized several limitations of the priorart blade designs. For example, the inventors have recognized thatconventional ceramic casting cores 64 are prone to breakage at theradial extremities, such as along the channel portion 68 where suddenchanges in stiffness occur. Furthermore, the inventors have recognizedthat the cooling system 19 does not distribute the cooling fluid acrossthe radial dimension of the cooling channel 18 in a manner consistentwith the heat transfer from the combustion fluid flow 25 across theradial dimension of the turbine blade 10. Heat load is typically amaximum proximate to a mid-span region of the airfoil. The ceramic coreis prone to breakage in the channel portion 68. To improve the core 64strength, the channel 22 is enlarged in a manner which causes adisproportionate amount of cooling flow to enter the cooling channel 18through the channel 22, where the heat load is relatively low, therebyproviding a mismatch between the heat load and the cooling capacity.

Thus, the present inventors have developed a blade with an improvedcooling arrangement which provides improved cooling performance while atthe same time being castable with a ceramic core which is less prone tobreakage at the radial extremities. The flow of cooling fluid across theradial dimension of the turbine blade in the present invention is moreclosely matched to the heat load distribution from the combustion fluidflow across the radial dimension of the turbine blade exterior.

In order to address the shortcomings of the conventional core design,the present inventors recognized that the radial extremities are proneto breakage, due to a low radius of curvature of the surfaces of thecore in those regions, which result in sudden changes in stiffness and aconcentration of stress. The present inventors recognized that in animproved core design, the radial extremities would be reshaped toincrease the radius of curvature of the radial extremity surfaces inorder to reduce or eliminate the sudden changes in stiffness in thosecritical regions.

Additionally, the present inventors recognized that the radius ofcurvature of the channel portion of the core, which shapes the radius ofcurvature of the channel of the turbine blade during the castingprocess, should be increased. With the improved core design, theincreased radius of curvature of the channel portion results in anextension of the channel of the turbine blade from the airfoil section,across the flow path line, through the platform section and into theroot section.

By increasing the radius of curvature of the channel of the turbineblade, a cross-sectional flow area through the entrance to the trailingedge impingement cooling channel increases. To maintain a desired totalcross-sectional area, in addition to increasing the radius of curvatureof the channel portion, the core design is further modified to move arib opening toward the root in the radial direction, which correspondsto lowering of a rib of the turbine blade in the radial direction, suchthat the increase in the cross-sectional flow area in the radialdirection caused by an increase in the radius of curvature of thechannel is offset to maintain the predetermined flow area. Additionally,to further improve the cooling system of the turbine blade such that thecooling effectiveness of cooling fluid at the inner and outer diameterregions is less than at an intermediate region between the inner andouter diameter regions, the core design is modified such that the widthof the trailing edge impingement cooling channel in the turbine blade isgreater at the intermediate region than at the inner and outer diameterregions. The modified width profile of the trailing edge impingementcooling channel more closely matches the heat exchange capacity of thecooling fluid to the heat load imposed across the radial extent of theblade.

FIGS. 4, 5 and 7 illustrate a turbine blade 110 in accordance with oneembodiment of the invention, including a leading edge 111, a trailingedge 113, a pressure side 115 and a suction side 117. FIG. 5 illustratesthat the turbine blade 110 includes an airfoil section 112, a rootsection 114, and a platform section 116 positioned between the airfoilsection 112 and the root section 114. The turbine blade 110 includes acooling system 119 positioned between the pressure side 115 and thesuction side 117, and includes several channels to pass a cooling fluidwhich is delivered from the root section 114. The cooling system 119includes a serpentine network of channels 144,146,148 including a firstchannel 144 which extends along the leading edge 111 from an innerdiameter region 128 to an outer diameter region 131 of the airfoilsection 112. At the outer diameter region 131, a first turn (not shown)is provided, to join the first channel 144 to a second channel 146,which extends from the outer diameter region 131 to the inner diameterregion 128. At the inner diameter region 128, a second turn 147 joinsthe second channel 146 with a third channel 148. A hot combustion fluidflow 125 from a combustor (not shown) of a turbine engine passes overthe exterior of the airfoil section 112, above a flow path line 141. Theflow path line 141 defines the radial extent of the area of the blade110 which is exposed to the hot combustion fluid flow 125. Thetemperature distribution of the hot combustion fluid flow 125 in aradial direction over the exterior of the airfoil section 112 is greaterin radial regions between the inner diameter region 128 and the outerdiameter region 131 than at the inner diameter region 128 and the outerdiameter region 131. In an exemplary embodiment, the temperaturedistribution of the hot combustion gas 125 is a maximum at anintermediate region 130 positioned half-way between the inner diameterregion 128 and the outer diameter region 131.

Cooling fluid is provided to the serpentine network of channels144,146,148 from a supply channel 160 within the root section 114, andpasses into the first channel 144 at the inner diameter region 128. Thecooling fluid subsequently passes to the first turn (not shown) at theouter diameter region 131, after which the cooling fluid passes into thesecond channel 146 and then flows from the outer diameter region 131 tothe second turn 147 at the inner diameter region 128, after which thecooling fluid passes to the third channel 148. As the cooling fluidpasses through the third channel 148, the cooling fluid partially passesthrough the impingement orifices 150 positioned between segmented ribs156 aligned between the third channel 148 and the trailing edge pin bankcooling channel 118. The cooling fluid subsequently passes throughimpingement orifices 152,154 of respectively spaced apart ribs 158,160,before exiting the trailing edge 113 through the orifices 162.

In addition to the serpentine network of channels 144,146,148, coolingfluid may enter the trailing edge pin bank cooling channel 118 bypassing from a cooling air supply channel 120 within the root section114, into a channel 122. The channel 122 begins within the root section114 and passes across the flow path line 141 before communicating withthe trailing edge pin bank cooling channel 118 in the airfoil section112. The cooling fluid within the trailing edge bin bank cooling channel118 passes through the orifices 152 of the segmented ribs 158, afterwhich the cooling fluid subsequently passes through the orifices 154 ofthe segmented ribs 160, and finally passes through the orifices 162 ofthe trailing edge 113, to exit the airfoil section 112. The turbineblade 110 and the airfoil section 112 illustrated in FIGS. 4-5 and 7 areexemplary, and the turbine blade and airfoil section may have variousalternative designs, with varying numbers of channels and/or structuraldesign, and still be within the scope of the embodiments of the presentinvention.

As illustrated in FIG. 5, the channel 122 interconnects the cooling airsupply channel 120 to the trailing edge pin bank cooling channel 118.The channel 122 receives the cooling fluid from the cooling air supplychannel 120 within the root section 114 and passes the cooling fluidthrough an entrance of the trailing edge pin bank cooling channel 118within the airfoil section 112. As illustrated in FIG. 5, a surface 127of the turbine blade 110 adjacent to the inner diameter region 128 isshaped with a radius of curvature 126. The channel 122 is shaped, basedon the surface 127, and thus the channel 122 is similarly shaped withthe radius of curvature 126 that defines a curvature of the channel 122.FIG. 6 illustrates a core 164 used to form the turbine blade 110 duringa casting process, as appreciated by one of skill in the art. The radiusof curvature 126 of the channel 122 and the surface 127 are definedbased on the radius of curvature 126 of a corresponding channel portion168 of the core 164 during the casting process. Thus, the channel 122extends from the cooling air supply channel 120 in the root section 114,through the platform section 116, and into the trailing edge pin bankcooling channel 118 in the airfoil section 112, based on a radius ofcurvature 126 of the channel portion 168 of the core 164. Alternatively,one may describe the cooling channel 118 as extending via a radius ofcurvature to below a flow path line of the blade. This iscounterintuitive to prior art blade designs where it was known toterminate a trailing edge cooling channel at or above the flow pathline, since there is no heat load being supplied into the blade materialbelow the flow path line.

FIG. 3 illustrates the conventional core 64 used to form theconventional turbine blade 10 of FIGS. 1-2, including a channel portion68 with the radius of curvature 26 that is used to shape the channel 22in the conventional turbine blade 10. As previously discussed, thechannel portion 68 of the conventional core 64 is prone to breakage, dueto the low radius of curvature 26 which results in sudden changes instiffness across the channel portion 68 and thus concentrated stress inthis region. As illustrated in FIG. 6, the design of the core 164 usedto form the turbine blade 110 features notable variations from thedesign of the conventional core 64 used to form the conventional turbineblade 10. As illustrated in FIGS. 3 and 6, the radius of curvature 126of the channel portion 168 of the core 164 is greater than the radius ofcurvature 26 of the channel portion 68 for the conventional core 64 (foran equivalent core design), to reduce or eliminate sudden changes instiffness across the channel portion 168, and thus locations ofconcentrated stress. Thus, the core 164 is less prone to breakage acrossthe channel portion 168, as compared to the core 64. Additionally, theincreased radius of curvature 126 of the channel portion 168 of the core164 will increase the radius of curvature 126 of the channel 122 of theturbine blade 110, such that the channel 122 extends from the coolingair supply channel 120 in the root section 114 to the trailing edge pinbank cooling channel 118 in the airfoil section 112. As illustrated inFIGS. 1-2, the channel 22 of the conventional turbine blade 10, shapedby the channel portion 68 of the conventional core 64, extends withinthe airfoil section 12 from the third channel 48 to the cooling channel18, and does not extend below the flow path line 41 or extend into theroot section 14.

As previously discussed, in order to improve the cooling system 119 ofthe turbine blade 110, the flow of cooling fluid at the inner and outerdiameter regions 128,131 is reduced below the flow of cooling fluid atthe intermediate region 130 between the inner and outer diameter regions128,131. Also, as previously discussed, the core 164 is designed, suchthat the radius of curvature 126 of the channel 122 is increased, whichconsequently increases the entrance to the cooling channel 118, in aradial direction, thereby increasing the flow of cooling fluid throughthe entrance to the cooling channel 118 at the inner diameter region128. In order to offset this increased flow, the increased entrance tothe cooling channel 118 in the radial direction is offset, such that theflow area through the entrance to the cooling channel 118 is mitigatedor limited to a predetermined flow area, thereby mitigating a flow ofcooling fluid through the entrance to the cooling channel 118 at theinner diameter region 128. To effect this offset, the rib opening 170 ofthe core 164 is lowered (relative to the rib opening 70 of theconventional core 64), such that the rib 138 of the turbine blade 110 isradially positioned to cooperate with the radius of curvature 126 of thechannel 122, to offset the increased entrance to the cooling channel 118in the radial direction, and maintain the predetermined cross-sectionalflow area through the channel 122 at the entrance of the trailing edgepin bank cooling channel 118. By increasing the radius of curvature 126of the channel portion 168 in the core 164 (from the conventional core64 design), the curvature of the channel 122 is reduced, to extend thechannel 122 from the airfoil section 112 through the platform section116 and into the root section 114, and subsequently increase thecross-sectional flow area through the channel 122 in the radialdirection at the entrance to the cooling channel 118. If the rib 138 isnot radially lowered in the airfoil section 112 design, to offset theincrease in the cross-sectional flow area in the radial direction, theinner diameter region 128 of the airfoil section 112 would be overcooledby an excessive cooling fluid flow through the channel 122 at theentrance to the cooling channel 118. Thus, by designing the core 164such that the rib opening 170 is appropriately positioned, thecross-sectional flow area through the channel 122 at the entrance to thecooling channel 118 is maintained at a predetermined flow rate, toenhance a cooling efficiency of the cooling system 119 of the turbineblade 110.

As previously discussed, FIG. 8 illustrates the trailing edge coolingchannel 18 of the conventional turbine blade 10, in which the width 32(measured between the pressure side 15 and the suction side 17) at theinner diameter region 28 is greater than the width 34 at theintermediate diameter region 30 positioned between the inner and outerdiameter regions 28,31. Thus, the flow of cooling fluid through theinner diameter region 28 of the conventional turbine blade 10 exceedsthe flow of cooling fluid flow through the intermediate diameter region30. The efficiency of the cooling system 19 arrangement is limited, asthe temperature distribution of the hot combustion fluid flow 25 isgreater at the intermediate region 30 than at the inner and outerdiameter regions 28,31, resulting in an overcooled inner and outerdiameter regions 28,31 and an undercooled intermediate region 30. Thecore 164 used to form the turbine blade 110 is designed, to furtherimprove the cooling system 119 of the turbine blade 110, to avoid theseshortcomings of the cooling system 19 of the conventional turbine blade10. The channel portion 172 of the core 164 is used to form the trailingedge cooling channel 118 of the turbine blade 110. The channel portion172 of the core 164 is reshaped, to adjust the width of the trailingedge pin bank cooling channel 118 (FIG. 7), which extends in a radialdirection from the inner diameter region 128 attached to the platformsection 116, to the outer diameter region 131. As further illustrated inFIG. 7, the channel portion 172 of the core 164 is reshaped, such that awidth 132,136 of the trailing edge pin bank cooling channel 118 at theinner diameter region 128 and at the outer diameter region 131 is lessthan the width 134 of the trailing edge pin bank cooling channel 118 ata mid-span region or an intermediate region 130 positioned between theinner diameter region 128 and the outer diameter region 131. In anexemplary embodiment, the intermediate region 130 is positioned at amidpoint or half-way between the inner diameter region 128 and the outerdiameter region 131. Although FIG. 7 illustrates that the maximum widthis the width 134 of the trailing edge pin bank cooling channel 118 atthe intermediate region 130 positioned half-way between the inner andouter diameter regions 128, 131, the embodiments of the presentinvention is not limited to this arrangement, provided that the width ofthe trailing edge pin bank cooling channel 118 is greater at a regionbetween the inner and outer diameter regions 128,131 than the width132,136 at the respective inner and outer diameter regions 128,131. Theredesigned core 164 and channel portion 172 are used to cast the coolingchannel 118 of the turbine blade 110 which conforms to the distributionof the heat temperature of the combustion fluid flow 125 over theexterior of the airfoil section 112, during the operation of the turbineengine. By adjusting the width 132,136 of the cooling channel 118 at theinner and outer diameter regions 128,131 to be less than the width 134of the cooling channel 118 at the intermediate region 130, the flow ofcooling fluid through the inner and outer diameter regions 128,131 ismitigated, thereby increasing the flow of cooling fluid at theintermediate region 130. Similarly, as previously discussed, the designof the core 164 is configured to shape the entrance to the coolingchannel 118 in a radial dimension, to also mitigate the flow of coolingfluid through the inner diameter region 128 of the cooling channel 118.The increased radius of curvature 126 of the channel portion 168 of thecore 164 increased the cross-sectional flow area (in the radialdimension) through the entrance to the cooling channel 118. However, thecore 164 design also featured lowering of the rib opening 170, such thatthe increased cross-sectional flow area (in the radial dimension) isoffset through the entrance to the cooling channel 118, to mitigate theflow of cooling fluid through the inner diameter region 128.

Thus, the core 164 design enhances the cooling efficiency of the coolingsystem 119, as the core 164 design is configured to shape the coolingchannel 118 such that the radial dimension of the entrance to thecooling channel 118 is mitigated or limited, to mitigate the flow ofcooling fluid through the inner diameter region 128 of the coolingchannel (and consequently that the flow of cooling fluid through theintermediate region 130 is increased). Similarly, the core 164 designenhances the cooling efficiency of the cooling system 119, as the core164 design shapes the cooling channel 118, in a lateral dimension, sothat the width 132,134 at the inner and outer diameter regions 128,131are less than the width 136 at the intermediate region 130, to mitigatethe flow of cooling fluid at the inner and outer diameter regions128,131 and enhance the flow of cooling fluid at the intermediate region130. Thus, the design of the core 164, and the trailing edge pin bankcooling channel 118 enhances the cooling efficiency of the trailing edgepin bank cooling channel 118 and the cooling system 119, by enhancingthe flow of cooling fluid at the intermediate region 130 and mitigatingthe flow of cooling fluid at the inner and outer diameter regions128,131. Although FIG. 7 illustrates a design to be used with regard tothe trailing edge pin bank cooling channel 118, an equivalent channeldesign may be utilized with regard to the other channels 144,146,148 ofthe cooling system 119 or in channels used in cooling systems other thanthe cooling system 119, for example.

Although FIG. 5 illustrates an embodiment of a turbine blade 110 designin which the channel 122 is shaped to extend from the cooling air supplychannel 120 in the root section 114 to the trailing edge pin bankcooling channel 118, and FIG. 7 illustrates an embodiment of a trailingedge pin bank cooling channel 118 design, the embodiments of FIGS. 5 and7 need not be incorporated into the same turbine blade 110. Thus, theturbine blade may be designed with the channel 122 design of FIG. 5, butwithout the trailing edge pin bank cooling channel 118 design of FIG. 7.Similarly, the turbine blade may be designed with the trailing edge pinbank cooling channel 118 design of FIG. 7, but without the channel 122design of FIG. 5.

While various embodiments of the present invention have been shown anddescribed herein, it will be obvious that such embodiments are providedby way of example only. Numerous variations, changes and substitutionsmay be made without departing from the invention herein. Accordingly, itis intended that the invention be limited only by the spirit and scopeof the appended claims.

1. A turbine blade including a trailing edge cooling channel wherein theimprovement comprises the trailing edge cooling channel having a maximumwidth proximate a mid-span region of an airfoil section of the blade andextending via a radius of curvature to below a flow path line of theblade.
 2. A turbine blade including an airfoil section, a root section,and a platform section between the airfoil section and the root section,wherein an exterior surface of said turbine blade is exposed to a hotcombustion gas above a flow path line, said turbine blade comprising: atrailing edge pin bank cooling channel in the airfoil section; a coolingair supply channel in the root section; a channel interconnecting thecooling air supply channel to the trailing edge pin bank coolingchannel, said channel having a radius of curvature to extend the channelfrom the cooling air supply channel in the root section, across the flowpath line and to the trailing edge pin bank cooling channel in theairfoil section.
 3. The turbine blade of claim 2, wherein said airfoilsection extends in a radial direction from an inner diameter regionattached to the platform section, to an outer diameter region; andwherein a width of the trailing edge pin bank cooling channel at theinner diameter region and at the outer diameter region is less than awidth of the trailing edge pin bank cooling channel at an intermediateregion positioned between the inner diameter region and the outerdiameter region.
 4. The turbine blade of claim 3, wherein said width ofthe trailing edge pin bank cooling channel is a maximum width at theintermediate region positioned half-way between the inner diameterregion and the outer diameter region.
 5. The turbine blade of claim 2,further comprising a rib being radially positioned to cooperate with theradius of curvature of the channel to define a predeterminedcross-sectional flow area through the channel at an entrance of thetrailing edge pin bank cooling channel.
 6. A ceramic core used to castthe turbine blade of claim 2 and including a portion defining thechannel.
 7. A turbine blade including an airfoil section, a root sectionand a platform section between the airfoil section and the root section,said turbine blade comprising: a trailing edge pin bank cooling channelin the airfoil section; wherein said airfoil section extends in a radialdirection from an inner diameter region attached to the platformsection, to an outer diameter region; and wherein a width of thetrailing edge pin bank cooling channel at the inner diameter region andat the outer diameter region is less than the width of the trailing edgepin bank cooling channel at an intermediate region positioned betweenthe inner diameter region and the outer diameter region.
 8. The turbineblade of claim 7, wherein an exterior surface of said turbine blade isexposed to a hot combustion gas above a flow path line, said turbineblade further comprising: a cooling air supply channel in the rootsection; and a channel to interconnect the cooling air supply channel tothe trailing edge pin bank cooling channel, said channel having a radiusof curvature to extend the channel from the cooling air supply channelin the root section across the flow path line and to the trailing edgepin bank cooling channel in the airfoil section.
 9. The turbine blade ofclaim 7, wherein said width of the trailing edge pin bank coolingchannel is a maximum width at the intermediate region positionedhalf-way between the inner diameter region and the outer diameterregion.
 10. The turbine blade of claim 8, further comprising a rib beingradially positioned to cooperate with the radius of curvature of thechannel to define a predetermined cross-sectional flow area through thechannel at an entrance of the trailing edge pin bank cooling channel.11. A ceramic core used to cast the turbine blade of claim 7.